High tip speed gas turbine engine

ABSTRACT

A gas turbine engine includes a compressor section and a turbine section. The turbine section includes a drive turbine and is located downstream of the compressor section. The gas turbine engine also includes a fan mechanically coupled to and rotatable with the drive turbine such that the fan is rotatable by the drive turbine at the same rotational speed as the drive turbine, the fan defining a fan pressure ratio and including a plurality of fan blades, each fan blade defining a fan tip speed. During operation of the gas turbine engine at a rated speed, the fan pressure ratio of the fan is less than 1.5 and the fan tip speed of each of the fan blades is greater than 1,250 feet per second.

FIELD

The present subject matter relates generally to a gas turbine engineconfigured to operate with a relatively high fan tip speed andrelatively low fan pressure ratio, and a method for operating the same.

BACKGROUND

A gas turbine engine generally includes a fan and a core arranged inflow communication with one another. Additionally, the core of the gasturbine engine generally includes, in serial flow order, a compressorsection, a combustion section, a turbine section, and an exhaustsection. In operation, air is provided from the fan to an inlet of thecompressor section where one or more axial compressors progressivelycompress the air until it reaches the combustion section. Fuel is mixedwith the compressed air using one or more fuel nozzles within thecombustion section and burned to provide combustion gases. Thecombustion gases are routed from the combustion section to the turbinesection. The flow of combustion gasses through the turbine sectiondrives the turbine section and is then routed through the exhaustsection, e.g., to atmosphere.

Typical gas turbine engines include a drive turbine within the turbinesection that is configured to drive, e.g., a low pressure compressor ofthe compressor section and the fan. In order to operate the gas turbineengine more efficiently, it is desirable to operate the drive turbine ata relatively high rotational speed. However, rotation of the fan atrelatively high rotational speeds can lead to inefficiencies, suchinefficiencies stemming from, e.g., shock losses and flow separation ofan airflow over fan blades of the fan.

Accordingly, certain gas turbine engines have been developed withreduction gearboxes that allow the fan to rotate slower than the driveturbine. However, certain gearboxes may add complication, weight, andexpense to the gas turbine engine. Therefore, a gas turbine engineconfigured to allow the drive turbine to operate at relatively high andefficient rotational speeds, while minimizing correspondinginefficiencies with the fan would be useful.

BRIEF DESCRIPTION

Aspects and advantages of the invention will be set forth in part in thefollowing description, or may be obvious from the description, or may belearned through practice of the invention.

In one exemplary embodiment of the present disclosure, a gas turbineengine is provided. The gas turbine engine includes a compressor sectionand a turbine section. The turbine section includes a drive turbine andis located downstream of the compressor section. The gas turbine enginealso includes a fan mechanically coupled to and rotatable with the driveturbine such that the fan is rotatable by the drive turbine at the samerotational speed as the drive turbine, the fan defining a fan pressureratio and including a plurality of fan blades, each fan blade defining afan tip speed. During operation of the gas turbine engine at a ratedspeed, the fan pressure ratio of the fan is less than 1.5 and the fantip speed of each of the fan blades is greater than 1,250 feet persecond.

In certain exemplary aspects, during operation of the gas turbine engineat the rated speed the fan pressure ratio is between 1.15 and 1.5.

In certain exemplary aspects, during operation of the gas turbine engineat the rated speed the fan pressure ratio is between 1.25 and 1.4.

In certain exemplary aspects, during operation of the gas turbine engineat the rated speed the fan tip speed of each of the fan blades isbetween about 1,350 feet per second and about 2,200 feet per second.

In certain exemplary aspects, during operation of the gas turbine engineat the rated speed the fan tip speed of each of the fan blades isgreater than about 1,450 feet per second.

In certain exemplary aspects, during operation of the gas turbine engineat the rated speed the fan tip speed of each of the fan blades isgreater than about 1.550 feet per second.

In certain exemplary aspects, the drive turbine of the turbine sectionis a low pressure turbine, wherein the gas turbine engine furtherincludes a high pressure turbine located upstream of the low pressureturbine. For example, in such an exemplary aspect the gas turbine enginemay further include a compressor section including a low pressurecompressor and a high pressure compressor, wherein the low pressurecompressor is driven by the low pressure turbine and the high pressurecompressor is driven by the high pressure turbine.

In certain exemplary aspects, the gas turbine engine may further includea nacelle surrounding and at least partially enclosing the fan. Forexample, in such an exemplary aspect, the gas turbine engine may furtherinclude an inlet pre-swirl feature located upstream of the plurality offan blades of the fan, the inlet pre-swirl feature attached to orintegrated into the nacelle.

In an exemplary aspect of the present disclosure, a method is providedof operating a direct drive gas turbine engine including a turbinesection with a drive turbine and a fan driven by the drive turbine. Themethod includes rotating the fan of the gas turbine engine with thedrive turbine of the turbine section of the gas turbine engine such thatthe fan rotates at an equal rotational speed as the drive turbine. Withsuch an exemplary aspect, the fan defines a fan pressure ratio less than1.5, and a fan blade of the fan defines a fan tip speed greater than1,250 feet per second.

In certain exemplary aspects, rotating the fan of the gas turbine engineof the drive turbine includes operating the gas turbine engine at arated speed.

In certain exemplary aspects, rotating the fan of the gas turbine enginewith the drive turbine includes rotating the fan of the gas turbineengine with the drive turbine such that the fan defines a fan pressureratio between 1.15 and 1.5.

In certain exemplary aspects, rotating the fan of the gas turbine enginewith the drive turbine includes rotating the fan of the gas turbineengine with the drive turbine such that the fan defines a fan pressureratio between 1.25 and 1.5.

In certain exemplary aspects, rotating the fan of the gas turbine enginewith the drive turbine includes rotating the fan of the gas turbineengine with the drive turbine such that the fan blade of the fan definesa fan tip speed between about 1,350 feet per second and about 2,200 feetper second.

In certain exemplary aspects, rotating the fan of the gas turbine enginewith the drive turbine includes rotating the fan of the gas turbineengine with the drive turbine such that the fan blade of the fan definesa fan tip speed greater than about 1,450 feet per second.

In certain exemplary aspects, rotating the fan of the gas turbine enginewith the drive turbine includes rotating the fan of the gas turbineengine with the drive turbine such that the fan blade of the fan definesa fan tip speed greater than about 1,550 feet per second.

In certain exemplary aspects, the method may further includepre-swirling a flow of air provided to the fan of the gas turbine engineduring operation of the gas turbine engine. For example, in such anexemplary aspect, pre-swirling the flow of air provided to the fan ofthe gas turbine engine includes pre-swirling the flow of air provided tothe fan of the gas turbine engine using an inlet pre-swirl featurelocated upstream of the fan blade of the fan and attached to orintegrated into a nacelle of the gas turbine engine.

In certain exemplary aspects, the drive turbine of the turbine sectionof the direct drive gas turbine engine is a low pressure turbine,wherein the turbine section of the direct drive gas turbine enginefurther includes a high pressure turbine, and wherein the direct drivegas turbine engine further includes a compressor section having a lowpressure compressor and a high pressure compressor and a nacellesurrounding and at least partially enclosing the fan.

These and other features, aspects and advantages of the presentinvention will become better understood with reference to the followingdescription and appended claims. The accompanying drawings, which areincorporated in and constitute a part of this specification, illustrateembodiments of the invention and, together with the description, serveto explain the principles of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

A full and enabling disclosure of the present invention, including thebest mode thereof, directed to one of ordinary skill in the art, is setforth in the specification, which makes reference to the appendedfigures, in which:

FIG. 1 is a schematic cross-sectional view of an exemplary gas turbineengine according to various embodiments of the present subject matter.

FIG. 2 is a close-up, schematic, cross-sectional view of a forward endof the exemplary gas turbine engine of FIG. 1.

FIG. 3 is a schematic view of an inlet to the exemplary gas turbineengine of FIG. 1, along an axial direction of the gas turbine engine ofFIG. 1.

FIG. 4 it is a schematic view of an inlet to a gas turbine engine inaccordance with another exemplary embodiment of the present disclosure.

FIG. 5 is a cross-sectional view of a part span inlet guide vane of theexemplary gas turbine engine of FIG. 1 at a first location along a spanof the part span inlet guide vane.

FIG. 6 is a cross-sectional view of the part span inlet guide vane ofthe exemplary gas turbine engine of FIG. 1 at a second location alongthe span of the part span inlet guide vane.

FIG. 7 is a close-up, schematic, cross-sectional view of a forward endof a gas turbine engine in accordance with another exemplary embodimentof the present disclosure.

FIG. 8 is a close-up, schematic, cross-sectional view of a forward endof a gas turbine engine in accordance with yet another exemplaryembodiment of the present disclosure.

FIG. 9 is a close-up, schematic, cross-sectional view of a forward endof a gas turbine engine in accordance with still another exemplaryembodiment of the present disclosure.

FIG. 10 is a close-up, schematic, cross-sectional view of a forward endof a gas turbine engine in accordance with yet another exemplaryembodiment of the present disclosure.

FIG. 10A is a close-up view of a trailing edge of a part span inletguide vane in accordance with an exemplary embodiment of the presentdisclosure.

FIG. 11 is a close-up, schematic, cross-sectional view of a forward endof a gas turbine engine in accordance with still another exemplaryembodiment of the present disclosure.

FIG. 12 is a close-up, schematic, cross-sectional view of a forward endof a gas turbine engine in accordance with yet another exemplaryembodiment of the present disclosure.

FIG. 13 is a schematic view of an inlet to the exemplary gas turbineengine of FIG. 12, along an axial direction of the gas turbine engine ofFIG. 12.

FIG. 14 it is a schematic view of an inlet to a gas turbine engine inaccordance with another exemplary embodiment of the present disclosure.

FIG. 15 is a perspective view of a pre-swirl contour of the exemplarygas turbine engine of FIG. 12.

FIG. 16 is a side view of the exemplary pre-swirl contour of theexemplary gas turbine engine of FIG. 12.

FIG. 17 is a lengthwise cross-sectional view of a plurality of pre-swirlcontours, including the exemplary pre-swirl contour of the exemplary gasturbine engine of FIG. 12.

FIG. 18 is a top view of a plurality of pre-swirl contours, includingthe exemplary pre-swirl contour of the exemplary gas turbine engine ofFIG. 12.

FIG. 19 is a close-up, schematic, cross-sectional view of a forward endof a gas turbine engine in accordance with still another exemplaryembodiment of the present disclosure.

FIG. 20 is a close-up, schematic, cross-sectional view of a forward endof a gas turbine engine in accordance with yet another exemplaryembodiment of the present disclosure.

FIG. 21 is a schematic, cross-sectional view of an outer nacelle of theexemplary gas turbine engine of FIG. 19, as viewed along Line 21-21 ofFIG. 19.

FIG. 22 is a close-up view of an airflow nozzle of the exemplary turbineengine of FIG. 21.

FIG. 23 is a close-up view of an airflow nozzle in accordance withanother exemplary embodiment of the present disclosure.

FIG. 24 is a close-up view of an airflow nozzle in accordance with yetanother exemplary embodiment of the present disclosure.

FIG. 25 is a close-up, cross-sectional view of the exemplary airflownozzle of FIG. 22.

FIG. 26 is a schematic, cross-sectional view of an outer nacelle and aportion of an airflow distribution system of a gas turbine engineaccordance with another exemplary embodiment of the present disclosure.

FIG. 27 is a perspective view of a plurality of swirl features of theexemplary airflow distribution system of FIG. 26 in accordance with anexemplary embodiment of the present disclosure.

FIG. 28 is a flow diagram depicting a method for operating a gas turbineengine in accordance with an exemplary aspect of the present disclosure.

FIG. 29 is a flow diagram depicting a method for operating a gas turbineengine in accordance with another exemplary aspect of the presentdisclosure.

DETAILED DESCRIPTION

Reference will now be made in detail to present embodiments of theinvention, one or more examples of which are illustrated in theaccompanying drawings. The detailed description uses numerical andletter designations to refer to features in the drawings. Like orsimilar designations in the drawings and description have been used torefer to like or similar parts of the invention.

As used herein, the terms “first”, “second”, and “third” may be usedinterchangeably to distinguish one component from another and are notintended to signify location or importance of the individual components.

The terms “forward” and “aft” refer to relative positions within a gasturbine engine, with forward referring to a position closer to an engineinlet and aft referring to a position closer to an engine nozzle orexhaust.

The terms “upstream” and “downstream” refer to the relative directionwith respect to fluid flow in a fluid pathway. For example, “upstream”refers to the direction from which the fluid flows, and “downstream”refers to the direction to which the fluid flows.

The singular forms “a”, “an”, and “the” include plural references unlessthe context clearly dictates otherwise.

Approximating language, as used herein throughout the specification andclaims, is applied to modify any quantitative representation that couldpermissibly vary without resulting in a change in the basic function towhich it is related. Accordingly, a value modified by a term or terms,such as “about”, “approximately”, and “substantially”, are not to belimited to the precise value specified. In at least some instances, theapproximating language may correspond to the precision of an instrumentfor measuring the value, or the precision of the methods or machines forconstructing or manufacturing the components and/or systems. Forexample, in certain contexts, the approximating language may refer tobeing within a 10% margin.

Here and throughout the specification and claims, range limitations maybe combined and interchanged, such that ranges identified include allthe sub-ranges contained therein unless context or language indicatesotherwise.

Referring now to the drawings, wherein identical numerals indicate thesame elements throughout the figures, FIG. 1 is a schematiccross-sectional view of a gas turbine engine in accordance with anexemplary embodiment of the present disclosure. More particularly, forthe embodiment of FIG. 1, the gas turbine engine is a high-bypassturbofan jet engine 10, referred to herein as “turbofan engine 10.” Asshown in FIG. 1, the turbofan engine 10 defines an axial direction A(extending parallel to a longitudinal centerline 12 provided forreference), a radial direction R, and a circumferential direction (i.e.,a direction extending about the axial direction A; see, e.g., FIG. 3).In general, the turbofan 10 includes a fan section 14 and a turbomachine16 disposed downstream from the fan section 14.

The exemplary turbomachine 16 depicted generally includes asubstantially tubular outer casing 18 that defines an annular inlet 20.The outer casing 18 encases, in serial flow relationship, a compressorsection including a booster or low pressure (LP) compressor 22 and ahigh pressure (HP) compressor 24; a combustion section 26; a turbinesection including a high pressure (HP) turbine 28 and a low pressure(LP) turbine 30; and a jet exhaust nozzle section 32. A high pressure(HP) shaft or spool 34 drivingly connects the HP turbine 28 to the HPcompressor 24. A low pressure (LP) shaft or spool 36 drivingly connectsthe LP turbine 30 to the LP compressor 22. The LP turbine 30 may also bereferred to as a “drive turbine”.

For the embodiment depicted, the fan section 14 includes a variablepitch fan 38 having a plurality of fan blades 40 coupled to a disk 42 ina spaced apart manner. More specifically, for the embodiment depicted,the fan section 14 includes a single stage fan 38, housing a singlestage of fan blades 40. As depicted, the fan blades 40 extend outwardlyfrom disk 42 generally along the radial direction R. Each fan blade 40is rotatable relative to the disk 42 about a pitch axis P by virtue ofthe fan blades 40 being operatively coupled to a suitable actuationmember 44 configured to collectively vary the pitch of the fan blades 40in unison. The fan 38 is mechanically coupled to and rotatable with theLP turbine 30, or drive turbine. More specifically, the fan blades 40,disk 42, and actuation member 44 are together rotatable about thelongitudinal axis 12 by LP shaft 36 in a “direct drive” configuration.Accordingly, the fan 38 is coupled with the LP turbine 30 in a mannersuch that the fan 38 is rotatable by the LP turbine 30 at the samerotational speed as the LP turbine 30.

Further, it will be appreciated that the fan 38 defines a fan pressureratio and the plurality of fan blades 40 each define a fan tip speed. Aswill be described in greater detail below, the exemplary turbofan engine10 depicted defines a relatively high fan tip speed and relatively lowfan pressure ratio during operation of the turbofan engine at a ratedspeed. As used herein, the “fan pressure ratio” refers to a ratio of apressure immediately downstream of the plurality of fan blades 40 duringoperation of the fan 38 to a pressure immediately upstream of theplurality of fan blades 40 during the operation of the fan 38. Also asused herein, the “fan tip speed” defined by the plurality of fan blades40 refers to a linear speed of an outer tip of a fan blade 40 along theradial direction R during operation of the fan 38. Further, still, asused herein, the term “rated speed” refers to a maximum operating speedof the turbofan engine 10, in which the turbofan engine 10 generates amaximum amount of power.

Referring still to the exemplary embodiment of FIG. 1, the disk 42 iscovered by rotatable front hub 48 aerodynamically contoured to promotean airflow through the plurality of fan blades 40. Additionally, theexemplary fan section 14 includes an annular fan casing or outer nacelle50 that circumferentially surrounds the plurality of fan blades 40 ofthe fan 38 and/or at least a portion of the turbomachine 16. Morespecifically, the nacelle 50 includes an inner wall 52 and a downstreamsection 54 of the inner wall 52 of the nacelle 50 extends over an outerportion of the turbomachine 16 so as to define a bypass airflow passage56 therebetween. Additionally, for the embodiment depicted, the nacelle50 is supported relative to the turbomachine 16 by a plurality ofcircumferentially spaced outlet guide vanes 55.

During operation of the turbofan engine 10, a volume of air 58 entersthe turbofan 10 through an associated inlet 60 of the nacelle 50 and/orfan section 14. As the volume of air 58 passes across the fan blades 40,a first portion of the air 58 as indicated by arrows 62 is directed orrouted into the bypass airflow passage 56 and a second portion of theair 58 as indicated by arrow 64 is directed or routed into the LPcompressor 22. The ratio between the first portion of air 62 and thesecond portion of air 64 is commonly known as a bypass ratio. For theembodiment depicted, the bypass ratio may generally be between about 7:1and about 20:1, such as between about 10:1 and about 18:1. The pressureof the second portion of air 64 is then increased as it is routedthrough the high pressure (HP) compressor 24 and into the combustionsection 26, where it is mixed with fuel and burned to provide combustiongases 66.

The combustion gases 66 are routed through the HP turbine 28 where aportion of thermal and/or kinetic energy from the combustion gases 66 isextracted via sequential stages of HP turbine stator vanes 68 that arecoupled to the outer casing 18 and HP turbine rotor blades 70 that arecoupled to the HP shaft or spool 34, thus causing the HP shaft or spool34 to rotate, thereby supporting operation of the HP compressor 24. Thecombustion gases 66 are then routed through the LP turbine 30 where asecond portion of thermal and kinetic energy is extracted from thecombustion gases 66 via sequential stages of LP turbine stator vanes 72that are coupled to the outer casing 18 and LP turbine rotor blades 74that are coupled to the LP shaft or spool 36, thus causing the LP shaftor spool 36 to rotate, thereby supporting operation of the LP compressor22 and/or rotation of the fan 38.

The combustion gases 66 are subsequently routed through the jet exhaustnozzle section 32 of the turbomachine 16 to provide propulsive thrust.Simultaneously, the pressure of the first portion of air 62 issubstantially increased as the first portion of air 62 is routed throughthe bypass airflow passage 56 before it is exhausted from a fan nozzleexhaust section 76 of the turbofan 10, also providing propulsive thrust.The HP turbine 28, the LP turbine 30, and the jet exhaust nozzle section32 at least partially define a hot gas path 78 for routing thecombustion gases 66 through the turbomachine 16.

It should be appreciated, however, that the exemplary turbofan engine 10depicted in FIG. 1 and described above is by way of example only, andthat in other exemplary embodiments, the turbofan engine 10 may have anyother suitable configuration. For example, in other exemplaryembodiments, the turbomachine 16 may include any other suitable numberof compressors, turbines, and/or shaft or spools. Additionally, theturbofan engine 10 may not include each of the features describedherein, or alternatively, may include one or more features not describedherein. For example, in other exemplary embodiments, the fan 38 may notbe a variable pitch fan. Additionally, although described as a“turbofan” gas turbine engine, in other embodiments the gas turbineengine may instead be configured as any other suitable ducted gasturbine engine.

Referring still to FIG. 1, and as previously discussed, the exemplaryturbofan engine 10 depicted in FIG. 1 is configured as a direct driveturbofan engine 10. In order to increase an efficiency of theturbomachine 16, the LP turbine 30 is configured to rotate at arelatively high rotational speed. Given the direct-drive configuration,such also causes the plurality of fan blades 40 of the fan 38 to rotateat a relatively high rotational speed. For example, during operation ofthe turbofan engine 10 at the rated speed, the fan tip speed of each ofthe plurality of fan blades 40 is greater than 1,250 feet per second.For example, in certain exemplary embodiments, during operation of theturbofan engine 10 at the rated speed, the fan tip speed of each of theplurality of fan blades 40 may be greater than about 1,350 feet persecond, such as greater than about 1,450 feet per second, such asgreater than about 1,550 feet per second such as up to about 2,200 feetper second.

Despite these relatively fan tip speeds, the fan 38 is, neverthelessdesigned to define a relatively low fan pressure ratio. For example,during operation of the turbofan engine 10 at the rated speed, the fanpressure ratio of the fan 38 is less than 1.5. For example, duringoperation of the turbofan engine 10 at the rated speed, the fan pressureratio may be between about 1.15 and about 1.5, such as between about1.25 and about 1.4.

As will be appreciated, operating the direct drive turbofan engine 10 insuch a manner may ordinarily lead to efficiency penalties of the fan 38due to shock losses and flow separation of an airflow over the fanblades 40, especially at the radially outer tips of the plurality of fanblades 40 of the fan 38. Accordingly, as will be described in muchgreater detail below, the turbofan engine 10 may further include one ormore inlet pre-swirl features upstream of the plurality of fan blades 40of the fan 38 to offset or minimize such efficiency penalties of the fan38. With the inclusion of such inlet pre-swirl features, the efficiencygains of the turbomachine 16 due to, e.g., increased rotational speedsof the LP turbine 30, outweigh the above identified potential efficiencypenalties.

Referring now also to FIG. 2, a close-up, cross-sectional view of thefan section 14 and forward end of the turbomachine 16 of the exemplaryturbofan engine 10 of FIG. 1 is provided. As stated, the turbofan engine10 includes an inlet pre-swirl feature located upstream of the pluralityof fan blades 40 of the fan 38 and attached to or integrated into thenacelle 50. More specifically, for the embodiment of FIGS. 1 and 2, theinlet pre-swirl feature is configured as a plurality of part span inletguide vanes 100. The plurality of part span inlet guide vanes 100 areeach cantilevered from of the outer nacelle 50 (such as from the innerwall 52 of the outer nacelle 50) at a location forward of the pluralityof fan blades 40 of the fan 38 along the axial direction A and aft ofthe inlet 60 of the nacelle 50. More specifically, each of the pluralityof part span inlet guide vanes 100 define an outer end 102 along theradial direction R, and are attached to/connected to the outer nacelle50 at the radially outer end 102 through a suitable connection means(not shown). For example, each of the plurality of part span inlet guidevanes 100 may be bolted to the inner wall 52 of the outer nacelle 50 atthe outer end 104, welded to the inner wall 52 of the outer nacelle 50at the outer end 102, or attached to the outer nacelle 50 in any othersuitable manner at the outer end 102.

Further, for the embodiment depicted, the plurality of part span inletguide vanes 100 extend generally along the radial direction R from theouter end 102 to an inner end 104 (i.e., an inner end 104 along theradial direction R). Moreover, as will be appreciated, for theembodiment depicted, each of the plurality of part span inlet guidevanes 100 are unconnected with an adjacent part span inlet guide vane100 at the respective inner ends 104 (i.e., adjacent part span inletguide vanes 100 do not contact one another at the radially inner ends104, and do not include any intermediate connection members at theradially inner ends 104, such as a connection ring, strut, etc.). Morespecifically, for the embodiment depicted, each part span inlet guidevane 100 is completely supported by a connection to the outer nacelle 50at the respective outer end 102 (and not through any structureextending, e.g., between adjacent part span inlet guide vanes 100 at alocation inward of the outer end 102 along the radial direction R). Aswill be discussed below, such may reduce an amount of turbulencegenerated by the part span inlet guide vanes 100.

Moreover, is depicted, each of the plurality of part span inlet guidevanes 100 do not extend completely between the outer nacelle 50 and,e.g., the hub 48 of the turbofan engine 10. More specifically, for theembodiment depicted, each of the plurality of inlet guide vane define anIGV span 106 along the radial direction R, and further each of theplurality of part span inlet guide vanes 100 further define a leadingedge 108 and a trailing edge 110. The IGV span 106 refers to a measurealong the radial direction R between the outer end 102 and the inner end104 of the part span inlet guide vane 100 at the leading edge 108 of thepart span inlet guide vane 100. Similarly, it will be appreciated, thatthe plurality of fan blades 40 of the fan 38 define a fan blade span 112along the radial direction R. More specifically, each of the pluralityof fan blades 40 of the fan 38 also defines a leading edge 114 and atrailing edge 116, and the IGV span 106 refers to a measure along theradial direction R between a radially outer tip and a base of the fanblade 40 at the leading edge 114 of the respective fan blade 40.

For the embodiment depicted, the IGV span 106 is at least about fivepercent of the fan blade span 112 and up to about fifty-five percent ofthe fan blade span 112. For example, in certain exemplary embodiments,the IGV span 106 may be between about fifteen percent of the fan bladespan 112 and about forty-five percent of the fan blade span 112, such asbetween about thirty percent of the fan blade span 112 and about fortypercent of the fan blade span 112.

Reference will now also be made to FIG. 3, providing an axial view ofthe inlet 60 to the turbofan engine 10 of FIGS. 1 and 2. As will beappreciated, for the embodiment depicted, the plurality of part spaninlet guide vanes 100 of the turbofan engine 10 includes a relativelylarge number of part span inlet guide vanes 100. More specifically, forthe embodiment depicted, the plurality of part span inlet guide vanes100 includes between about twenty part span inlet guide vanes 100 andabout fifty part span inlet guide vanes 100. More specifically, for theembodiment depicted, the plurality of part span inlet guide vanes 100includes between about thirty part span inlet guide vanes 100 and aboutforty-five part span inlet guide vanes 100, and more specifically,still, the embodiment depicted includes thirty-two part span inlet guidevanes 100. Additionally, for the embodiment depicted, each of theplurality of part span inlet guide vanes 100 are spaced substantiallyevenly along the circumferential direction C. More specifically, each ofthe plurality of part span inlet guide vanes 100 defines acircumferential spacing 118 with an adjacent part span inlet guide vane100, with the circumferential spacing 118 being substantially equalbetween each adjacent part span inlet guide vane 100.

Although not depicted, in certain exemplary embodiments, the number ofpart span inlet guide vanes 100 may be substantially equal to the numberof fan blades 40 of the fan 38 of the turbofan engine 10. In otherembodiments, however, the number of part span inlet guide vanes 100 maybe greater than the number of fan blades 40 of the fan 38 of theturbofan engine 10, or alternatively, may be less than the number of fanblades 40 of the fan 38 of the turbofan engine 10.

Further, should be appreciated, that in other exemplary embodiments, theturbofan engine 10 may include any other suitable number of part spaninlet guide vanes 100 and/or circumferential spacing 118 of the partspan inlet guide vanes 100. For example, referring now briefly to FIG.4, an axial view of an inlet 60 to a turbofan engine 10 in accordancewith another exemplary embodiment of the present disclosure is provided.For the embodiment of FIG. 4, the turbofan engine 10 includes less thantwenty part span inlet guide vanes 100. More specifically, for theembodiment of FIG. 4, the turbofan engine 10 includes at least eightpart span inlet guide vanes 100, or more specifically includes exactlyeight part span inlet guide vanes 100. Additionally, for the embodimentof FIG. 4, the plurality of part span inlet guide vanes 100 are notsubstantially evenly spaced along the circumferential direction C. Forexample, at least certain of the plurality of part span inlet guidevanes 100 define a first circumferential spacing 118A, while other ofthe plurality of part span inlet guide vanes 100 define a secondcircumferential spacing 118B. For the embodiment depicted, the firstcircumferential spacing 118A is at least about twenty percent greaterthan the second circumferential spacing 118B, such as at least abouttwenty-five percent greater such as at least about thirty percentgreater, such as up to about two hundred percent greater. Notably, aswill be described in greater detail below, the circumferential spacing118 refers to a mean circumferential spacing between adjacent part spaninlet guide vanes 100. The non-uniform circumferential spacing may,e.g., offset structure upstream of the part span inlet guide vanes 100.

Referring now back to the embodiment of FIG. 2, it will be appreciatedthat each of the plurality of part span inlet guide vanes 100 isconfigured to pre-swirl an airflow 58 provided through the inlet 60 ofthe nacelle 50, upstream of the plurality of fan blades 40 of the fan38. As briefly discussed above, pre-swirling the airflow 58 providedthrough the inlet 60 of the nacelle 50 prior to such airflow 58 reachingthe plurality of fan blades 40 of the fan 38 may reduce separationlosses and/or shock losses, allowing the fan 38 to operate with therelatively high fan tip speeds described above with less losses of inefficiency.

For example, referring first to FIG. 5, a cross-sectional view of onepart span inlet guide vane 100 along the span of the part span inletguide vanes 100, as indicated by Line 5-5 in FIG. 2, is provided. As isdepicted, the part span inlet guide vane 100 is configured generally asan airfoil having a pressure side 120 and an opposite suction side 122,and extending between the leading edge 108 and the trailing edge 110along a camber line 124. Additionally, the part span inlet guide vane100 defines a chord line 126 extending directly from the leading edge108 to the trailing edge 110. The chord line 126 defines an angle ofattack 128 with an airflow direction 129 of the airflow 58 through theinlet 60 of the nacelle 50. Notably, for the embodiment depicted, theairflow direction 129 is substantially parallel to the axial direction Aof the turbofan engine 10. For the embodiment depicted, the angle ofattack 128 at the location depicted along the span 106 of the part spaninlet guide vanes 100 is at least about five degrees and up to aboutthirty-five degrees. For example, in certain embodiments, the angle ofattack 128 at the location depicted along the span 106 of the part spaninlet guide vane 100 may be between about ten degrees and about thirtydegrees, such as between about fifteen degrees and about twenty-fivedegrees.

Additionally, the part span inlet guide vane 100, at the locationdepicted along the span 106 of the part span inlet guide vane 100defines a local swirl angle 130 at the trailing edge 110. The “swirlangle” at the trailing edge 110 of the part span inlet guide vane 100,as used herein, refers to an angle between the airflow direction 129 ofthe airflow 58 through the inlet 60 of the nacelle 50 and a referenceline 132 defined by a trailing edge section of the pressure side 120 ofthe part span inlet guide vane 100. More specifically, the referenceline 132 is defined by the aft twenty percent of the pressure side 120,as measured along the chord line 126. Notably, when the aft twentypercent the pressure side 120 defines a curve, the reference line 132may be straight-line average fit of such curve (e.g., using least meansquares).

Further, it will be appreciated, that a maximum swirl angle 130 refersto the highest swirl angle 130 along the span 106 of the part span inletguide vane 100. For the embodiment depicted, the maximum swirl angle 130is defined proximate the radially outer end 102 of the part span inletguide vane 100 (e.g., at the outer ten percent of the span 106 of thepart span inlet guide vanes 100), as is represented by the cross-sectiondepicted in FIG. 5. For the embodiment depicted, the maximum swirl angle130 of each part span inlet guide vane 100 at the trailing edge 110 isbetween five degrees and thirty-five degrees. For example, in certainexemplary embodiments, the maximum swirl angle 130 of each part spaninlet guide vane 100 at the trailing edge 110 may be between twelvedegrees and twenty-five degrees.

Moreover, it should be appreciated that for the embodiment of FIG. 2,the local swirl angle 130 increases from the radially inner end 104 tothe radially outer end 102 of each part span inlet guide vane 100. Forexample, referring now also to FIG. 6, a cross-sectional view of a partspan inlet guide vane 100 at a location radially inward from thecross-section viewed in FIG. 5, as indicated by Line 6-6 in FIG. 2, isprovided. As is depicted in FIG. 6, and as stated above, the part spaninlet guide vane 100 defines the pressure side 120, the suction side122, the leading edge 108, the trailing edge 110, the camber line 124,and chord line 126. Further, the angle of attack 128 defined by thechord line 126 and the airflow direction 129 of the airflow 58 throughthe inlet 60 of the nacelle 50 at the location along the span 106depicted in FIG. 6 is less than the angle of attack 128 at the locationalong the span 106 depicted in FIG. 5 (e.g., may be at least abouttwenty percent less, such as at least about fifty percent less, such asup to about one hundred percent less). Additionally, the part span inletguide vane 100 defines a local swirl angle 130 at the trailing edge 110at the location along the span 106 of the part span inlet guide vane 100proximate the inner end 104, as depicted in FIG. 6. As stated above, thelocal swirl angle 130 increases from the radially inner end 104 to theradially outer end 102 of each part span inlet guide vanes 100.Accordingly, the local swirl angle 130 proximate the outer end 102 (seeFIG. 5) is greater than the local swirl angle 130 proximate the radiallyinner end 104 (see FIG. 6; e.g., the radially inner ten percent of thespan 106). For example, the local swirl angle 130 may approach zerodegrees (e.g., may be less than about five degrees, such as less thanabout two degrees) at the radially inner end 104.

Notably, including part span inlet guide vanes 100 of such aconfiguration may reduce an amount of turbulence at the radially innerend 104 of each respective part span inlet guide vane 100. Additionally,such a configuration may provide a desired amount of pre-swirl at theradially outer ends of the plurality of fan blades 40 of the fan 38(where the speed of the fan blades 40 is the greatest) to provide adesired reduction in flow separation and/or shock losses that mayotherwise occur due to a relatively high speed of the plurality of fanblades 40 at the fan tips during operation of the turbofan engine 10.

Referring generally to FIGS. 2, 3, 5, and 6, it will be appreciated thatfor the embodiment depicted, the plurality of part span inlet guidevanes 100 further define a solidity. The solidity is defined generallyas a ratio of a chord length (i.e., a length of the chord line 126) ofeach part span inlet guide vane 100 to a circumferential spacing 118 ofthe plurality of part span inlet guide vanes 100. More specifically, forthe purposes of defining the solidity, the circumferential spacing 118refers to the mean circumferential spacing 118 calculated using thefollowing equation:

2×n×r _(m) ² ÷n _(b)  (Equation 1);

wherein r_(m) is the mean radius of the plurality of part span inletguide vanes 100 and n_(b) is the number of part span inlet guide vanes100. The mean radius, r_(m), may refer to a position halfway along theIGV span 106, relative to the longitudinal centerline 12 of the turbofanengine 10. Notably, for the purposes of calculating solidity, the chordlength refers to the chord length at the mean radius, r_(m). For theembodiment depicted, the solidity is between about 0.5 and is about 1.5.For example, in certain exemplary embodiments, the solidity of the partspan inlet guide vanes 100 may be between about 0.7 and 1.2, such asbetween about 0.9 and about 1.0. Such a configuration may ensure desiredamount of pre-swirl during operation of the turbofan engine 10.

Notably, the plurality of part span inlet guide vanes 100 depicted inFIGS. 1 through 6 are generally configured to pre-swirl a portion of anairflow through the inlet 60 of the outer nacelle 50 in a rotationaldirection that is the same as a rotational direction of the plurality offan blades 40 of the fan 38. For example, for the exemplary embodimentof FIGS. 1 through 6, the plurality of fan blades 40 of the fan 38 areconfigured to rotate clockwise when viewed forward looking aft and theplurality of part-span inlet guide vanes 100 (and other pre-swirlfeatures discussed herein) are configured to pre-swirl a portion of theairflow through the inlet 60 of the outer nacelle 50 in the samedirection. However, in other exemplary embodiments the gas turbineengine may include a fan 38 with fan blades 40 configured to rotatecounter-clockwise when viewed forward looking aft, in which case theplurality of part-span inlet guide vanes 100 (or other pre-swirlfeatures discussed herein) may instead be mirrored such that they areconfigured to pre-swirl airflow in an opposite rotational direction thanthe direction depicted. Further, in still other exemplary embodiments,the plurality of part-span inlet guide vanes 100 (or other pre-swirlfeatures discussed herein) may be configured to pre-swirl an airflow inan opposite rotational direction as the plurality of fan blades 40 ofthe fan 38.

Additionally, it should be appreciated that the exemplary part spaninlet guide vanes 100 depicted in FIGS. 1 through 6 are provided by wayof example only. In other exemplary embodiments, the plurality of partspan inlet guide vanes 100 may have any other suitable configuration forproviding a desired amount of pre-swirl upstream of a plurality of fanblades 40 of a fan 38 of a gas turbine engine. For example, referringgenerally to FIGS. 7 through 11, part span inlet guide vanes 100 inaccordance with various other exemplary embodiments of the presentdisclosure are provided. Each of the exemplary turbofan engines 10 andexemplar) part span inlet guide vanes 100 of FIGS. 7 through 11 may beconfigured in substantially the same manner as the exemplary turbofanengine 10 and part span inlet guide vanes 100 described above withreference to, e.g., FIGS. 1 and 2.

For example, the exemplary turbofan engines 10 of FIGS. 7 through 11each generally include a turbomachine 16 and a fan section 14, anddefine an axial direction A, a radial direction R, and a circumferentialdirection C (i.e., a direction extending about the axial direction A;see, e.g., FIG. 3). The turbomachine 16, although not depicted, includesa turbine section having a drive turbine, or LP turbine 30 (see FIG. 1),mechanically coupled to a fan 38 of the fan section 14 through, for theembodiment depicted, an LP shaft 36. Additionally, the fan 38 includes aplurality of fan blades 40 rotatable about a longitudinal centerline 12of the turbomachine 16. The plurality of fan blades 40 of the fan 38 aresurrounded by, and enclosed by, an outer nacelle 50 of the turbofanengine 10, the outer nacelle 50 including an inner wall 52. In order toprovide for a pre-swirling of an airflow 58 through an inlet 60 of theouter nacelle 50, the turbofan engine 10 further includes a plurality ofpart span inlet guide vanes 100. As stated, the exemplary part spaninlet guide vanes 100 of FIGS. 7 through 11 are each configured in asimilar manner to the exemplary part span inlet guide vanes 100described above with reference to FIGS. 1 and 2. Accordingly, each ofthe plurality of part span inlet guide vanes 100 may be attached in acantilevered fashion to the inner wall 52 of the outer nacelle 50 at alocation forward of the plurality of fan blades 40 of the fan 38 of theturbofan engine 10 and aft of the inlet 60.

However, referring particularly to FIG. 7, for the embodiment depicted,the plurality of part span inlet guide vanes 100 are further configuredas variable part span inlet guide vanes. More specifically, each of theplurality of part span inlet guide vanes 100 depicted in FIG. 7 includea body portion 134 and a tail portion 136 (the tail portion 136 locatedaft of the body portion 134). Each of the body portion 134 and tailportion 136 extends substantially from a radially outer end 102 of thepart span inlet guide vanes 100 to a radially inner end 104 of the partspan inlet guide vanes 100. The body portion 134 may be substantiallyfixed, while the tail portion 136 may be configured to rotate about alongitudinal pivot axis 138 of the respective part span inlet guide vane100 by a motor 140 or other pitch change mechanism. For the embodimentdepicted, the longitudinal pivot axis 138 is substantially parallel tothe radial direction R, however in other embodiments, the longitudinalpivot axis 138 may extend in any other suitable direction (e.g., may be“swept”; see FIG. 9). Rotation of the tail portion 136 of the part spaninlet guide vanes 100 may effectively vary a swirl angle 130 of therespective part span inlet guide vane 100. Accordingly, with such anexemplary embodiment, the turbofan engine 10 may be configured toprovide minimal pre-swirl during certain operating conditions, andprovide maximum pre-swirl during other operating conditions. Forexample, in certain exemplary embodiments, the turbofan engine 10 may beconfigured to provide minimal pre-swirl when the fan 38 is rotating at arelatively slow rotational speed (such that the fan 38 defines arelatively low fan tip speed), and may further be configured to providea maximum pre-swirl when the fan 38 is rotating at a relatively highrotational speed (such that the fan 38 defines a relatively high fan tipspeed, such as during takeoff operating modes).

It should be appreciated, that the exemplary variable part span inletguide vanes 100 depicted in FIG. 7 are provided by way of example only.In other exemplary embodiments, any other suitable variable part spaninlet guide vanes 100 may be provided. For example, in other exemplaryembodiments, the tail portion 136 may not extend along an entire span106 of the part span inlet guide vanes 100, and instead may be limitedto, e.g., a radially outer half, or other portion, of the part spaninlet guide vanes 100. Additionally, any other suitable hardware may beprovided for varying a swirl angle 130 of the variable part span inletguide vanes 100. For example, in other embodiments, the tail portion 136may not rotate about the longitudinal pivot axis 138, and instead maytranslate, e.g., forward and aft to modify the swirl angle 130. Otherconfigurations are contemplated as well (e.g., pneumatic variable partspan inlet guide vanes 100 using air to vary an effective swirl angle).

Additionally, referring now particularly to FIG. 8, for the exemplaryembodiment depicted, the exemplary part span inlet guide vanes 100 eachdefine a variable span 106. More specifically, for the embodiment ofFIG. 8, the inner end 104 of each of the plurality of part span inletguide vanes 100 is movable generally along the radial direction Rbetween an extended position (depicted) and a retracted position(depicted in phantom). For example, for the embodiment depicted each ofthe plurality of part span inlet guide vanes 100 defines a first span106A when the extended position and a second span 106B when in theretracted position. The second span 106B may be between about twentypercent and about ninety percent of the first span 106A. For example,the second span 106B may be between about thirty percent and abouteighty percent of the first span 106A, such as between about fortypercent and about sixty percent of the first span 106A.

Referring still to FIG. 8, the exemplary part span inlet guide vane 100is formed generally of a base portion 142 and an extendable portion 144.The extendable portion 144 is at least partially nested within the baseportion 142 when the part span inlet guide vane 100 is in the retractedposition (depicted in phantom), and more specifically is substantiallycompletely nested within the base portion 142 when part span inlet guidevane 100 is in the retracted position. Additionally, as is depicted, theextendable portion 144 is positioned substantially outside the baseportion 142 when the part span inlet guide vane 100 is in the extendedposition.

For the embodiment of FIG. 8, the extendable portion 144 is movablegenerally along the radial direction R by an extension rod 146 operableby a motor 148. However, in other embodiments, any other suitableassembly may be provided for moving part span inlet guide vane 100between the extended position and the retracted position. Additionally,although the exemplary part span inlet guide vane 100 is depicted asincluding a single extendable portion 144, in other exemplaryembodiments, one or more of the plurality of part span inlet guide vanes100 may instead include a plurality of extendable portions 144 which maynest when moved to the retracted position (e.g., may include up to tenextendable portions 144). Further, other configurations for moving theradially inner end 104 generally along the radial direction R between anextended position and a retracted position are contemplated as well. Forexample, extendable portions 144 may fold or pivot to the “retracted”position.

Accordingly, with such an exemplary embodiment, the turbofan engine 10may be configured to provide minimal pre-swirl during certain operatingconditions (e.g., by moving the part span inlet guide vane 100 to theretracted position), and provide maximum pre-swirl during otheroperating conditions (e.g., by moving the part span inlet guide vane 100to the extended position). For example, in certain exemplaryembodiments, the turbofan engine 10 may be configured to provide minimalpre-swirl when the fan 38 is rotating at a relatively slow rotationalspeed (such that the fan 38 defines a relatively low fan tip speed), andmay further be configured to provide a maximum pre-swirl when the fan 38is rotating at a relatively high rotational speed (such that the fan 38defines a relatively high fan tip speed, such as during takeoffoperating modes).

Referring now particularly to FIG. 9, for the exemplary embodimentdepicted, the exemplary part span inlet guide vanes 100 are configuredas “swept” part span inlet guide vanes. More specifically, as isdepicted, the exemplary part span inlet guide vanes 100 each define alongitudinal axis 150 extending halfway between the leading edge 108 anda trailing edge 110 from the radially inner end 104 to the radiallyouter end 102. Additionally, the exemplary turbofan engine 10 defines areference plane 152, or more specifically, the radial direction R andcircumferential direction C the turbofan engine 10 together define areference plane 152. The longitudinal axis 150 of each of the pluralityof part span inlet guide vanes 100 intersects the reference plane 152and defines a sweep angle 154 with the reference plane 152. For theembodiment depicted, the sweep angle 154 with the reference plane 152 isgreater than about five degrees and up to about forty degrees. Morespecifically, for the embodiment depicted, the sweep angle 154 with thereference plane 152 is between about ten degrees and about thirtydegrees, such as between about fifteen degrees and about twenty-fivedegrees. Inclusion of part span inlet guide vanes 100 defining a sweepangle 154 in accordance the embodiment of FIG. 9 may provide certainacoustic benefits and/or pre-swirl benefits during operation of theturbofan engine 10.

Notably, although for the embodiment depicted the exemplary part spaninlet guide vanes 100 extend in a substantially straight direction fromthe radially inner end 104 to the radially outer end 102, in otherembodiments one or more of the plurality of part span inlet guide vanes100 may instead extend in a curved direction (i.e., a curved part-spaninlet guide vane 100). With such a configuration, for the purposes ofdefining the sweep angle 154, the longitudinal axis 150 of such a partspan inlet guide vane 100 may refer to a line extending between a pointhalfway between the leading edge 108 and trailing edge 110 at theradially outer end 102 of the part span inlet guide vane 100 to a pointhalfway between the leading edge 108 and the trailing edge 110 at theradially inner end 104 of the part span inlet guide vane 100.

Additionally, although for the embodiment depicted the plurality of partspan inlet guide vanes 100 are configured with a forward-to-aft sweep,in other exemplary embodiments of the present disclosure, the pluralityof part span inlet guide vanes 100 may instead define an aft-to-forwardsweep (i.e., the sweep angle 154 may be negative).

Referring now to FIG. 10, the exemplary part span inlet guide vanes 100are configured to include a trailing edge 110 that is sculpted. Morespecifically, for the embodiment of FIG. 10, the trailing edge 110 ofeach of the plurality of part span inlet guide vanes 100 defines anon-linear sculpted shape. As used herein, the term “non-linear sculptedshape”, with reference to a trailing edge 110, refers to any shapehaving sequential variations at the trailing edge 110 (i.e., sequentialincreases and decreases) in a local chord length and/or camber length(see, e.g., FIGS. 5 and 6, depicting cord line 126 in camber line 24)along the span 106 of the part span inlet guide vane 100. For example,referring particularly to Circle 10A in FIG. 10, providing a close-upview of a section of the trailing edge 110, the trailing edge 110defines a periodic, sinusoidal wave shape. The wave shape of thetrailing edge 110 defines a cycle distance 156 greater than about fivepercent of a span 106 of the part span inlet guide vane 100, and lessthan about thirty-three percent of the span 106 of the part span inletguide vane 100. Accordingly, the wave shape repeats at least three timesand up to about twenty times along the span 106 of the part span inletguide vane 100. Additionally, the wave shape defines a height 158 (i.e.,a peak-to-valley height) less than the cycle distance 156, such asbetween about five percent of the cycle distance 156 and about ninetypercent of the cycle distance 156.

It should be appreciated, however, that in other embodiments, theplurality of part span inlet guide vanes 100 may have any other suitablesculpting at the trailing edge 110. For example, in other embodiments,the height 158 may be equal to, or greater than, the cycle distance 156,such as up to about five times greater than the cycle distance 156.Additionally, in other embodiments, the same shape may not repeat, andfurther, the sculpting shape may include any other suitable shape inaddition to, or in the alternative to, waves. For example, the shape mayinclude triangles, other polygons, semicircles, etc.

Inclusion of a plurality of part span guide vanes 100 having thetrailing edge 110 in accordance with one or more of these embodimentsmay provide acoustic benefits to the turbofan engine 10 during operationof the turbofan engine 10 by, e.g., increasing a mixing of the wake ofthe part span inlet guide vanes 100 and a bulk airflow 58 through theinlet 60 of the nacelle 50.

Referring now to FIG. 11, the plurality of part span inlet guide vanes100 for the embodiment depicted are further configured to provide acompensation airflow 160 to a trailing edge 110 of the plurality of partspan inlet guide vanes 100 to minimize a wake of the part span inletguide vanes 100. More specifically, for the embodiment of FIG. 11, theturbofan engine 10 further includes a compensation air supply assembly162 in airflow communication with a high pressure air source. Thecompensation air supply assembly 162 generally includes a compensationair supply duct 164 defining an inlet 166 in airflow communication withthe high pressure air source, which for the embodiment depicted is thecompressor section of the turbofan 10. For example, the compensation airsupply duct 164 may be configured to receive bleed air from thecompressor section of the turbofan engine 10. Notably, however, in otherembodiments, the compensation air supply duct 164 may instead receivehigh pressure air from any other suitable high pressure air source. Forexample, in other exemplary embodiments, the high pressure air sourcemay instead be the bypass airflow duct 56 at a location downstream ofthe plurality of fan blades 40 of the fan 38. Additionally, in one ormore these embodiments, the compensation air supply assembly 162 mayfurther include an air compressor 166 (depicted in phantom) configuredto increase a pressure of the compensation airflow 160 through thecompensation air supply duct 164. Notably, although the supply duct 164is depicted as a single, continuous, and separate supply duct 164, inother embodiments, the composition air supply duct 164 may have anyother suitable configuration. For example, the duct 164 may be formed ofa plurality of sequential ducts, may be formed integrally with othercomponents of the turbofan engine 10, and/or may be split off into aplurality of parallel airflow ducts to provide compensation airflow 160to each of the plurality of part span inlet guide vanes 100 (similar tothe distribution an extension air tubes 210, 212 of the embodiment ofFIG. 21).

Further, the compensation air supply duct 164 extends through at leastone of the plurality of part span inlet guide vanes 100, and provides acavity 168 of the part span inlet guide vane 100 with the high pressurecomposition airflow 160. As is depicted, the each of the plurality ofpart span inlet guide vanes 100 for the embodiment depicted furtherdefines a trailing edge opening 170, which is in airflow communicationwith the cavity 168, and thus is in airflow communication with thecompensation air supply duct 164 of the compensation air supply assembly162. Accordingly, with such a configuration, the high pressurecomposition airflow 160 may be provided from the compensation air supplyassembly 162 to the cavity 168 of the part span inlet guide vane 100,and further through the trailing edge opening 170 of the part span inletguide vane 100 during operation of the turbofan engine 10 to reduce awake formed by the respective part span inlet guide vane 100.

It should be appreciated that although described as a “cavity” 168, inother embodiments the cavity 168 may be configured as any suitableopening or passage within the part span inlet guide vane 100 to allow aflow of air therethrough. Additionally, it should be appreciated that inother exemplary embodiments, the plurality of part span inlet guidevanes 100 may instead include any other suitable manner of pneumaticallyreducing the wake of the respective part span inlet guide vanes 100. Forexample, in other exemplary embodiments, the trailing edge opening 170of each part span inlet guide vane 100 may instead be configured as,e.g., a plurality of trailing edge of openings spaced, e.g., along aspan 106 of the respective part span inlet guide vane 100 at thetrailing edge 110.

It should further be appreciated that in still other embodiments of thepresent disclosure any other suitable inlet pre-swirl feature may beprovided at a location upstream of the plurality of fan blades 40 of thefan 38 of the gas turbine engine and downstream of an inlet 60 of anouter nacelle 50. For example, referring now to FIG. 12, an inletpre-swirl feature of a gas turbine engine in accordance with anotherexemplary embodiment of the present disclosure is provided. Morespecifically, FIG. 12 depicts a turbofan engine 10 in accordance with anembodiment of the present disclosure, configured in substantially thesame manner as the exemplary turbofan engine 10 described above withreference to FIGS. 1 and 2. Accordingly, the exemplary turbofan engine10 of FIG. 12 generally includes a turbomachine 16 and a fan section 14.The turbomachine 16, although not depicted, includes a turbine sectionhaving a drive turbine, or LP turbine 30 (see FIG. 1), mechanicallycoupled to a fan 38 of the fan section 14 through, for the embodimentdepicted, an LP shaft 36. Additionally, the fan 38 includes a pluralityof fan blades 40 rotatable about a longitudinal centerline 12 of theturbomachine 16. The plurality of fan blades 40 of the fan 38 aresurrounded by, and enclosed by, an outer nacelle 50 of the turbofanengine 10, the outer nacelle 50 including an inner wall 52. Further, theexemplary turbofan engine 10 includes an inlet pre-swirl featureattached to or integrated with the inner wall 52 of the outer nacelle 50at a location forward of the plurality of fan blades 40 of the fan 38.

However, for the embodiment of FIG. 12, the inlet pre-swirl feature doesnot include a plurality of part span inlet guide vanes 100, and insteadis configured as a plurality of pre-swirl contours 172 positionedforward of the fan blades 40 of the fan 38 along the axial direction Aand extending inwardly along the radial direction R. Each of theplurality of pre-swirl contours 172 may be spaced along thecircumferential direction C of the turbofan engine 10. For example,referring now also to FIG. 13, providing a schematic, axial view of theinlet 60 to the turbofan engine 10, each of the plurality of pre-swirlcontours 172 are spaced substantially evenly along the circumferentialdirection C, such that each adjacent pre-swirl contour 172 defines asubstantially uniform circumferential spacing 174. Additionally, itshould be appreciated that the exemplary turbofan engine 10 may includeany suitable number of pre-swirl contours 172. For example, in certainexemplary embodiments the plurality of pre-swirl contours 172 includesbetween about five pre-swir contours 172 and about eighty pre-swirlcontours 172, such as between about thirty pre-swirl contours 172 andabout fifty pre-swirl contours 172, and more specifically, for theembodiment depicted, includes thirty-two pre-swirl contours 172.

However, in other exemplary embodiments, the plurality of pre-swirlcontours 172 may have any other suitable spacing. For example, referringbriefly to FIG. 14, providing a schematic, axial view of an inlet 60 ofa turbofan engine 10 in accordance with another exemplary embodiment ofthe present disclosure, the plurality of pre-swirl contours 172 maydefine a non-uniform circumferential spacing 174. For example, at leastcertain of the plurality of pre-swirl contours 172 define a firstcircumferential spacing 174A, while other of the plurality of pre-swirlcontours 172 define a second circumferential spacing 174B. For theembodiment depicted, the first circumferential spacing 174A is at leastabout twenty percent greater than the second circumferential spacing174B, such as at least about twenty-five percent greater such as atleast about thirty percent greater, such as up to about two hundredpercent greater. The circumferential spacing 174 refers to a meancircumferential spacing between adjacent pre-swirl contours 172.

Referring now also to FIGS. 15 through 18, various other views of one ormore of the plurality pre-swirl contours 172 of FIG. 12 are provided.More specifically, FIG. 15 provides a perspective view of the exemplarypre-swirl contour 172 of FIG. 12; FIG. 16 provides a side view of theexemplary pre-swirl contours 172 are of FIG. 12; FIG. 17 provides across-sectional view of a plurality of pre-swirl contours 172, includingthe exemplary pre-swirl contour 172 of FIG. 12; and FIG. 18 provides atop view of a plurality of pre-swirl contours 172, including theexemplary pre-swirl contour 172 FIG. 12.

Referring first particularly to FIG. 15, it will be appreciated that forthe embodiment depicted, the plurality of pre-swirl contours 172 areformed integrally with the inner wall 52 of the outer nacelle 50 to forma monolithic component. For example, the inner wall 52 of the outernacelle 50 may be formed by casting to include the plurality ofpre-swirl contours 172, or alternatively the inner wall 52 of the outernacelle 50 may be stamped to include the plurality of pre-swirl contours172, or alternatively, still, the inner wall 52 of the outer nacelle 50may be formed using a suitable additive manufacturing technique.However, it should be appreciated that in other exemplary embodimentsthe plurality of pre-swirl contours 172 may instead be formed separatelyfrom the inner wall 52 and attached to the inner wall 52 of the outernacelle 50 (or some other component of the outer nacelle 50) in anyother suitable manner.

Referring particularly to FIGS. 16 and 17, it will be appreciated thateach of the plurality of pre-swirl contours 172 defines a height 176along the radial direction R, and further extends generally from an aftend 178 to a forward end 180. Additionally, for the exemplary embodimentdepicted, the pre-swirl contours 172 each define an arcuate shape alongthe axial direction A, extending from the forward end 180 to the aft end178. Accordingly, the height 176 of the pre-swirl contour 172 variesalong a length thereof. More particularly, at the forward ends 180 ofthe pre-swirl contours 172, the height 176 of each respective pre-swirlcontour 172 is approximately equal to zero (e.g., less than five percentof a maximum height 176), and similarly at the aft ends 178 of thepre-swirl contours 172 the height 176 of each respective pre-swirlcontours 172 is approximately equal to zero (e.g., less than fivepercent of a maximum height 176).

Moreover, referring now also particularly to FIG. 18, each of theplurality pre-swirl contours 172 further defines a ridge line 182, eachridge line 182 tracking a peak height 176 of the respective pre-swirlcontour 172 between the forward and aft ends 180, 178 of the respectivepre-swirl contour 172. A maximum height 176 of each of the plurality ofpre-swirl contours 172 for the embodiment depicted is located within amiddle seventy-five percent of the respective ridge line 182, asmeasured along a total length of the respective ridge line 182. Morespecifically, for the embodiment depicted, the maximum height 176 ofeach of the plurality of pre-swirl contours 172 is located within amiddle fifty percent of the respective ridge line 182.

Additionally, the maximum height 176 of each of the plurality ofpre-swirl contours 172 may be sufficient to provide a desired amount ofpre-swirl to an airflow 58 received through an inlet 60 of the outernacelle 50 (see FIG. 12). For example, in certain exemplary embodiments,the maximum height 176 of each of the plurality of pre-swirl contours172 may be between about two percent and about forty percent of a fanblade span 112 of a fan blade 40 of the fan 38 (see FIG. 12). Forexample, in certain exemplary embodiments, the maximum height 176 ofeach of the plurality of pre-swirl contours 172 may be between aboutfive percent and about thirty percent of a fan blade span 112 of a fanblade 40, such as between about ten percent and about twenty-fivepercent of a fan blade span 112 of a fan blade 40.

Further, the plurality of pre-swirl contours 172 define a swirl angle184. With reference to the pre-swirl contours 172, the swirl angle 184refers to an angle of the ridge line 182 relative to an airflowdirection 129 of the airflow 58 through the inlet 60 of the nacelle 50during operation of the turbofan engine 10, which may be parallel to theaxial direction A of the turbofan engine 10. Referring particularly toFIG. 18, a maximum swirl angle 184 is defined by the aft twenty-fivepercent of the ridge line 182. Additionally, for the embodimentdepicted, the maximum swirl angle 184 of each of the plurality ofcontours 172 is between about five degrees and about forty) degrees. Forexample, the maximum swirl angle 184 of each of the plurality ofcontours 172 may be between about ten degrees and about thirty degrees,such as between about fifteen degrees and about twenty-five degrees.Notably, although the aft twenty-five percent of the ridge line 182 forthe embodiment depicted is substantially straight, in other exemplaryembodiments, it may define a curve. In such embodiments, the maximumswirl angle 184 may be defined with a reference line equal to an averageof the aft twenty-five percent of the ridge line 182. The average of theaft twenty-five percent of the ridge line 182 may be found using. e.g.,least mean squares or suitable method.

It will be appreciated, however, that the exemplary pre-swirl contours172 described herein with reference to FIGS. 12 through 18 are providedby way of example only. In other exemplary embodiments, the plurality ofpre-swirl contours 172 may have any other suitable shape and/orconfiguration. For example, in other exemplary embodiments, one or moreof the plurality of pre-swirl contours 172 may not define an arcuateshape, and may, for example, define a height 176 greater than zero atone or both of the forward end 180 and aft end 178. Additionally, inother exemplary embodiments, the plurality of pre-swirl contours 172 maynot each define substantially the same shape. For example, in otherexemplary embodiments, one or more the plurality of pre-contours 172 maydefine a maximum height 176 greater than an adjacent pre-swirl contour172.

Additionally, it will be appreciated that inclusion of one or more ofthe plurality of pre-swirl contours 172 in accordance with an exemplaryembodiment of the present disclosure may provide for an increasedefficiency of the turbofan engine 10 when operating with, e.g.,relatively high fan tip speeds. For example, the plurality of pre-swirlcontours 172 may provide an amount of pre-swirl to an airflow 58 throughan inlet 60 of a nacelle 50 of the turbofan engine 10, such that theairflow 58 at the radially outer ends of the fan blades 40 of the fan 38is less susceptible to separation from the plurality of fan blades 40and/or shock losses.

It should further be appreciated that in still other embodiments of thepresent disclosure any other suitable inlet pre-swirl feature may beprovided located upstream of the plurality of fan blades 40 of the fan38 of the gas turbine engine. For example, referring now to FIG. 19, aninlet pre-swirl feature in accordance with yet another exemplaryembodiment of the present disclosure is provided. More specifically,FIG. 19 depicts a turbofan engine 10 in accordance with an embodiment ofthe present disclosure, configured in substantially the same manner asthe exemplary turbofan engine 10 described above with reference to FIGS.1 and 2. Accordingly, the exemplary turbofan engine 10 of FIG. 19generally includes a turbomachine 16 and a fan section 14. Theturbomachine 16 includes a compressor section and, although notdepicted, a turbine section having a drive turbine, or LP turbine 30(see FIG. 1), mechanically coupled to a fan 38 of the fan section 14through, for the embodiment depicted, an LP shaft 36. Additionally, thefan 38 includes a plurality of fan blades 40 rotatable about alongitudinal centerline 12 of the turbomachine 16. The plurality of fanblades 40 of the fan 38 are surrounded by, and enclosed by, an outernacelle 50 of the turbofan engine 10, the outer nacelle 50 including aninner wall 52. Downstream of the fan 38 of the fan section 14, the outernacelle 50 defines a bypass airflow passage 56 with the turbomachine 16.Further, the exemplary turbofan engine 10 includes an inlet pre-swirlfeature attached to or integrated with the inner wall 52 of the outernacelle 50 at a location forward of the plurality of fan blades 40 ofthe fan 38.

However, for the embodiment of FIG. 19, the inlet pre-swirl feature doesnot include a plurality of part span inlet guide vanes 100, and insteadis configured as an airflow delivery system 186. More specifically, forthe embodiment of FIG. 19 the inner wall 52 of the outer nacelle 50defines a plurality of openings 188 located forward of the plurality offan blades 40 of the fan 38 along the axial direction A. The inletpre-swirl feature includes these plurality of openings 188, with theplurality of openings 188 configured to provide a swirl airflow 190upstream of the plurality of fan blades 40 of the fan 38 at a swirlangle 192 greater than zero relative to the radial direction R of theturbofan engine 10 (and more specifically relative to a local referenceplane defined by the axial direction A and the radial direction R). Asis depicted, for the embodiment of FIG. 19 the airflow delivery system186 generally includes an air tube 194 extending between an inlet 196and an outlet 198. As will be discussed in greater detail below, theoutlet 198 of the air tube 194 is in airflow communication with theplurality of openings 188 defined by the inner wall 52 of the outernacelle 50. Additionally, the inlet 196 of the air tube 194 is inairflow communication with a high pressure air source for receiving theswirl airflow 190. For the embodiment depicted, the high pressure airsource is the bypass airflow passage 56 at a location downstream of theplurality of fan blades 40 the fan 38.

As is depicted in phantom, in certain embodiments, the airflow deliverysystem 186 may further include a door 200 (i.e., a door, scoop, or otherstructural component) to scoop air into the inlet 196 of the air tube194. The door 200 may be movable between an open position and closedposition depending on, e.g., an operating condition of the gas turbineengine. For example, the door 200 may move to the open position when itis desirable to provide pre-swirling of an airflow 58 through the inlet60 of the outer nacelle 50. As is also depicted in phantom, the airflowdelivery system 186 may further include an air compressor 202, the aircompressor 202 in airflow communication with the air tube 194. The aircompressor 202 may act to increase a pressure of the swirl airflow 190through the air tube 194 to increase an amount of, e.g., pre-swirlprovided by the inlet pre-swirl feature depicted.

Notably, however, in other exemplary embodiments any other suitable highpressure air source may be provided. For example, referring now to FIG.20, a cross-sectional view of a gas turbine engine in accordance withanother exemplary embodiment of the present disclosure is provided. Theexemplary gas turbine engine of FIG. 20 is configured in substantiallythe same manner as the exemplary turbofan engine 10 described above withreference to FIG. 19. However, for the embodiment of FIG. 20, the airtube 194 of the airflow delivery system 186 is in airflow communicationwith a different high pressure air source. More specifically, for theembodiment of FIG. 20, the high pressure air source is a compressor ofthe compressor section of the turbomachine 16. More specifically, still,for the embodiment of FIG. 20, the high pressure air source is acompressor bleed valve 204 of the compressor section of the turbofanengine 10. However, in still other exemplary embodiments, any othersuitable high pressure air source may be provided.

Referring back to FIG. 19, the turbofan engine 10, or rather the airflowdelivery system 186 of the turbofan engine 10, further includes aplurality of airflow nozzles 206, with each airflow nozzle 206positioned at one of the openings 188 defined by the inner wall 52 ofthe nacelle 50. Referring now also to FIG. 21, a cross-sectional view ofa section of the outer nacelle 50 defining the openings 188 andincluding the airflow nozzles 206 is provided along Line 21-21 of FIG.19. As is depicted, the air tube 194 of the airflow delivery system 186further includes a plurality of segments. For example, in the embodimentdepicted, the air tube 194 includes a supply air tube 208 in airflowcommunication with the inlet 196 for receiving the swirl airflow 190from the high pressure air source. Additionally, the air tube 194includes a distribution air tube 210 extending from the supply air tube208 and, for the embodiment depicted, in the circumferential direction Csubstantially three hundred sixty degrees within the outer nacelle 50.Further, the air tube 194 includes a plurality of extension air tubes212 extending between the distribution air tube 210 and the plurality ofairflow nozzles 206, with each of the extension air tubes 212 defining arespective outlet 198 of the air tube 194. Accordingly, with such anembodiment, the air tube 194 further defines a plurality of outlets 198.

As is depicted, the airflow nozzles 206 each define an airflow direction214, the airflow direction 214 being the direction in which the swirlairflow 190 is provided through the openings 188 of the inner wall 52 ofthe outer nacelle 50. In certain exemplary embodiments, the airflowdirection 214 of each of the respective airflow nozzles 206 may extendalong a centerline 215 of each of the respective airflow nozzles 206.Additionally, for the embodiment depicted the airflow direction 214defines the swirl angle 192. Accordingly, for the embodiment depicted,the swirl angle 192 may refer to an angle between the airflow direction214 of the plurality of airflow nozzles 206 and the radial direction Rof the turbofan engine 10, or more specifically, for the embodimentdepicted, the swirl angle 192 refers to an angle between the airflowdirection 214 and a reference plane defined by the radial direction Rand the axial direction A of the turbofan engine 10. In certainexemplary embodiments, the swirl angle 192 is between five degrees andthirty-five degrees. For example, in certain embodiments the swirl angle192 may be between ten degrees and thirty degrees, such as betweenfifteen degrees and twenty-five degrees.

Further, the plurality of airflow nozzles 206 may include any suitablenumber of airflow nozzles 206, such as between about five airflownozzles 206 and about one hundred airflow nozzles 206. Morespecifically, for the embodiment depicted, the plurality of airflownozzles 206 includes eight airflow nozzles 206. However, in otherembodiments, the turbofan engine 10 of FIG. 21 may include the samenumber of airflow nozzles 206 as, e.g., the exemplary turbofan engine 10described above with reference to FIGS. 1 through 3 includes part spaninlet guide vanes 100. For example, in certain exemplary embodiments,the turbofan engine 10 may include at least twenty airflow nozzles 206,such as at least thirty airflow nozzles 206, and up to about fiftyairflow nozzles 206, such as up to about forty-five airflow nozzles 206.

Referring now briefly to FIG. 22, providing a close-up view of one ofthe exemplary airflow nozzles 206, it will be appreciated that for theembodiment depicted, the plurality of airflow nozzles 206 are formedseparately from the inner wall 52 of the outer nacelle 50 and attachedto the inner wall 52 of the outer nacelle 50. Additionally, for theembodiment depicted, the plurality of airflow nozzles 206 each extendthrough a respective opening 188 in the inner wall 52 of the outernacelle 50. It should be appreciated, however, that in other exemplaryembodiments, any other suitable configuration of airflow nozzles 206 maybe provided. For example, referring briefly to FIG. 23, in otherexemplary embodiments, one or more of the plurality of airflow nozzles206 may be formed integrally with the inner wall 52 of the outer nacelle50 (e.g., by casting, stamping, additive manufacturing, etc.), andfurther, referring now briefly to FIG. 24, in other exemplaryembodiments, one or more of the plurality of airflow nozzles 206 may notextend through the opening 188 of the inner wall 52 of the outer nacelle50. Moreover, in still other exemplary embodiments, one or more of theplurality of airflow nozzles 206 may be flush with the opening 188defined in the inner wall 52 the outer nacelle 50, or alternatively, theturbofan engine 10, and more specifically, the airflow delivery system186, may not include airflow nozzles 206 altogether.

Notably, for the exemplary turbofan engine 10 described above withreference to, e.g., FIGS. 19 and 21, the airflow delivery system 186 isconfigured to provide the swirl airflow 190 generally in a directionaligned with a reference plane defined by the radial direction R andcircumferential direction C (i.e., the plane depicted in FIG. 21).However, in other exemplary embodiments, the airflow delivery system 186may instead be configured to provide the swirl airflow 190 at an anglegreater than zero with the reference plane defined by the radialdirection R and the circumferential direction C. For example, referringnow briefly to FIG. 25, providing a cross-sectional view of an airflownozzle 206 in accordance with another exemplary embodiment of thepresent disclosure, the airflow delivery system 186 may be configured toprovide the swirl airflow 190 at an angle 216 between. e.g., about fivedegrees and about fifty degrees, such as between about ten degrees andabout thirty-five degrees with the reference plane defined by thecircumferential direction C of the radial direction R. With theseembodiments, the airflow nozzles 206 may be referred to as “swept”airflow nozzles.

It should be appreciated, however, that in still other exemplaryembodiments, the airflow delivery system 186 of the turbofan engine 10he have any other suitable configuration. For example, referring nowalso to FIG. 26, a cross-sectional view of a section of an outer nacelle50 defining openings 188 in accordance with another exemplary embodimentof the present disclosure is provided. The cross-sectional view of FIG.6 may be the same view provided in FIGS. 21, taking along Line 21-21 ofFIG. 19.

Additionally, the embodiment of FIG. 26 may be similar to the exemplar)embodiment of FIG. 19, described above. For example, as is depicted, theouter nacelle 50 generally includes an inner wall 52 defining aplurality of openings 188 and the airflow delivery system 186 generallyincludes an air tube 194. The air tube 194 extends between an inlet 196and an outlet 198, the outlet 198 being in airflow communication withthe plurality of openings 188 and the inlet 196 being in airflowcommunication with a high pressure air source for receiving the swirlairflow 190 (see FIG. 19). Moreover, as with the exemplary embodiment ofFIG. 19, the air tube 194 generally includes a supply air tube 208 and adistribution air tube 210, the distribution air tube 210 extendinggenerally in a circumferential direction C within the outer nacelle 50.

However, for the embodiment of FIG. 26, instead of including a pluralityof extension air tube 212 (see FIG. 21), the airflow delivery system 186includes a plenum 218. The plenum 218 is generally configured as anannular plenum extending circumferentially within the outer nacelle 50around the openings 188 and between the distribution air tube 210 theinner wall 52 of the outer nacelle 50. Accordingly, for the embodimentdepicted, the plenum 218 is defined at least in part by the inner wall52 of the outer nacelle 50 and the distribution air tube 210, as well asa forward wall and an aft wall (not shown). However, in other exemplaryembodiments, the plenum 218 may be defined by any other suitablecomponents of, e.g., the outer nacelle 50 and/or the airflow deliverysystem 186.

Moreover, for the exemplary embodiment depicted, the airflowdistribution system 186 further includes a plurality of swirl featurespositioned within the plenum 218 for directing the swirl airflow 190through the plenum 218 to the openings 188. More specifically, referringnow also to FIG. 27, providing a perspective view of a portion of theplurality of swirl features of the airflow distribution system 186 ofFIG. 26, for the embodiment depicted, each of the plurality of swirlfeatures is configured as an airfoil 220 extending generally between thedistribution air tube 210 and the inner wall 52 of the nacelle 50. Aswill be appreciated, the plurality of airfoils 220 are configured toswirl the airflow 190 provided to the plenum 218 prior to such airflow190 being provided through the plurality of openings 188 in the innerwall 52 of the nacelle 50.

Further, as is depicted, each of the plurality of airfoils 220 generallydefines an airflow direction 222, the airflow direction 222 being thedirection in which the swirl airflow 190 is provided through theopenings 188 of the inner wall 52. For the embodiment depicted, theairflow direction 222 may be substantially equal to a direction of areference line 224 defined by a trailing edge of a pressure side 226 ofthe respective airfoil 220, the reference line 224 being defined by theaft twenty percent of the pressure side 226. More specifically, thereference line 224 is defined by the aft twenty percent of the pressureside 120, as measured along a chord line of the respective airfoil 220.Notably, when the aft twenty percent the pressure side 226 defines acurve, the reference line 224 may be straight-line average fit of suchcurve (e.g., using least mean squares).

Additionally, for the embodiment depicted the airflow direction 222 (andreference line 224) defines a swirl angle 192. Accordingly, for theembodiment depicted, the swirl angle 192 may refer to an angle betweenthe airflow direction 222 of a respective airfoil 220 and the radialdirection R of the turbofan engine 10, or more specifically, for theembodiment depicted, the swirl angle 192 refers to an angle between theairflow direction 222 and a reference plane defined by the radialdirection R and the axial direction A of the turbofan engine 10. Incertain exemplary embodiments, the swirl angle 192 is between fivedegrees and eighty-five degrees. For example, in certain embodiments theswirl angle 192 may be between ten degrees and eighty degrees, such asbetween thirty degrees and seventy-five degrees.

Further, the airflow distribution system 186 may include any suitablenumber of airfoils 220 within the plenum 218, such as between about fiveairfoils 220 and about one hundred airfoils 220. For example, in certainembodiments, airflow distribution system 186 of FIG. 26 may include thesame number of airfoils 220 as, e.g., the exemplary turbofan engine 10described above with reference to FIGS. 1 through 3 includes part spaninlet guide vanes 100. For example, in certain exemplary embodiments,the airflow distribution system 186 may include at least twenty airfoils220, such as at least thirty airfoils 220, and up to about fiftyairfoils 220, such as up to about forty-five airfoils 220.

Referring now to FIG. 28, a flow diagram is provided of a method 300 foroperating a direct drive gas turbine engine in accordance with anexemplary aspect of the present disclosure. The exemplary direct driveturbofan engine may be configured in accordance with one or more theexemplary gas turbine engines described above with reference to FIGS. 1through 27. Accordingly, for example, the direct drive gas turbineengine may include a turbine section having a drive turbine and a fansection having a fan driven by the drive turbine.

The exemplary method 300 generally includes at (302) rotating the fan ofthe gas turbine engine with the drive turbine of the turbine section ofthe gas turbine engine such that the fan rotates at an equal rotationalspeed as the drive turbine. Additionally, for the exemplary aspectdepicted, rotating the fan of the gas turbine engine with the driveturbine at (302) include at (304) rotating the fan of the gas turbineengine with the drive turbine such that the fan defines a fan pressureratio less than 1.5. More specifically, for the exemplary aspectdepicted, rotating the fan of the gas turbine engine at (304) furtherincludes at (306) rotating the fan of the gas turbine engine with thedrive turbine such that the fan defines a fan pressure ratio between1.15 and 1.5, and further still at (308) rotating the fan of the gasturbine engine with the drive turbine such that the fan defines a fanpressure ratio between 1.25 and 1.5.

Referring still to FIG. 28, rotating the fan of the gas turbine enginewith the drive turbine at (304) further includes at (310) rotating thefan of the gas turbine engine with the drive turbine such that a fanblade the fan defines a fan tip speed greater than 1,250 feet persecond. More specifically, for the exemplary aspect depicted, rotatingthe fan of the gas turbine engine with the drive turbine at (304)further includes at (312) rotating the fan of the gas turbine enginewith the drive turbine such that the fan blade of the fan defines a fantip speed between about 1,350 feet per second and about 2.200 feet persecond. More specifically, still, for the exemplary aspect depicted,rotating the fan of the gas turbine engine with the drive turbine at(304) further includes at (314) rotating the fan of the gas turbineengine with the drive turbine such that the fan blade of the fan definesa fan tip speed greater than about 1,450 feet per second, and at (316)rotating the fan of the gas turbine engine with the drive turbine suchthat the fan blade of the fan defines a fan tip speed greater than about1,550 feet per second.

Further, as is also depicted, for the embodiment FIG. 28, rotating thefan of the gas turbine engine with the drive turbine at (304) includesat (318) operating the gas turbine engine at a rated speed. For example,operating the gas turbine engine at the rated speed at (318) may includeoperating the gas turbine at a maximum speed to produce a maximum ratedpower.

Moreover, the exemplary method 300 further includes at (320)pre-swirling a flow of air provided to the fan of the gas turbine engineduring operation of the gas turbine engine. For the exemplary aspectdepicted, pre-swirling the flow of air at (320) includes at (322)pre-swirling the flow of air provided to the fan of the gas turbineengine using an inlet pre-swirl feature located upstream of theplurality of fan blades of the fan and attached to or integrated into anacelle of the gas turbine engine. In certain exemplary aspects, theinlet pre-swirl feature may be configured in accordance with one or moreof the exemplary inlet pre-swirl features described above with referenceto FIGS. 1 through 27. By way of example only, in certain exemplaryaspects, pre-swirling the flow of air at (322) may include one or moreof the steps (408) through (414) of the exemplary method 400 describedbelow. However, in other embodiments, any other suitable inlet pre-swirlfeature and/or method may be utilized.

Operating a direct drive gas turbine engine in accordance with theexemplary aspect described above with reference to FIG. 28 may result ina more efficiently operated gas turbine engine. Further, when theairflow provided to the fan is pre-swirled, such may reduce an amount ofseparation or shock losses of the airflow with the fan despite therelatively high fan tip speeds at which the fan is operated.

Referring now to FIG. 29, a flow diagram of a method 400 for operating adirect drive gas turbine engine in accordance with another exemplaryaspect of the present disclosure is provided. The exemplary method 400may be utilized with the exemplary gas turbine engines described abovewith reference to FIGS. 19 through 27. Accordingly, for example, thedirect drive gas turbine engine may include a turbomachine, a fansection, and an outer nacelle, with the turbomachine including a driveturbine and the fan section including a fan.

Similar to the exemplary method 300, the exemplary method 400 includesat (402) rotating the fan of the gas turbine engine with the driveturbine of the turbomachine such that the fan rotates at an equalrotational speed as the drive turbine. For the exemplary aspectdepicted, rotating the fan with the drive turbine at (402) includes at(404) rotating the fan of the gas turbine engine such that a fan bladeof the fan defines a fan tip speed greater than 1,250 feet per second.Additionally, rotating the fan of the drive turbine at (402) furtherincludes, for the exemplary aspect depicted, at (406) rotating the fanof the gas turbine engine such that the fan defines a fan pressure ratioless than 1.5.

Referring still to FIG. 29, the method further includes at (408)receiving a pre-swirl airflow from a high pressure air source and at(410) transferring the pre-swirl airflow received from the high pressureair source to a plurality of airflow nozzles positioned at an inner wallof the outer nacelle at a location forward of the fan of the fansection. In certain exemplary aspects, the high pressure air source maybe, e.g., a bypass airflow passage of the direct drive gas turbineengine, or a compressor section of the direct drive gas turbine engine.Additionally, transferring the pre-swirl airflow at (410) may include,e.g., transferring the pre-swirl airflow through one or more air tubesor ducts defined within the direct drive gas turbine engine.

Further, the exemplary method 400 includes at (412) providing thepre-swirl airflow at a pre-swirl angle through the inner wall of theouter nacelle at a location forward of the fan of the fan section. Forthe exemplary aspect depicted, providing the pre-swirl airflow at thepre-swirl angle through the inner wall of the outer nacelle at (412)includes at (414) providing the pre-swirl airflow through a plurality ofopenings defined by the inner wall of the outer nacelle. Morespecifically, for the exemplary aspect depicted, providing the pre-swirlairflow through the plurality of openings defined by the inner wall ofthe outer nacelle at (414) includes providing the pre-swirl airflowthrough the plurality of airflow nozzles, each of the plurality ofairflow nozzles positioned at, or in airflow communication with, arespective opening defined by the inner wall of the outer nacelle at alocation forward of the fan of the fan section. It should beappreciated, however, that in other exemplary aspects, the gas turbineengine may not include the airflow nozzles, and instead may include anyother suitable structure for providing the pre-swirl airflow through theplurality of openings at the pre-swirl angle at (414).

Further, for the exemplary aspect depicted, the pre-swirl angle at whichthe pre-swirl airflow is provided through the inner wall of the outernacelle is between about five degrees and about thirty-five degrees.Additionally, the pre-swirl angle may be defined relative to, e.g., aradial direction of the direct drive gas turbine engine, or morespecifically, relative to a plane defined by the radial direction and anaxial direction of the gas turbine engine.

Operating a direct drive gas turbine engine in accordance with theexemplary aspect described above with reference to FIG. 29 may result ina more efficiently operated gas turbine engine. Further, when theairflow provided to the fan is pre-swirled, such may reduce an amount ofseparation or shock losses of the airflow with the fan despite therelatively high fan tip speeds at which the fan is operated.

This written description uses examples to disclose the invention,including the best mode, and also to enable any person skilled in theart to practice the invention, including making and using any devices orsystems and performing any incorporated methods. The patentable scope ofthe invention is defined by the claims, and may include other examplesthat occur to those skilled in the art. Such other examples are intendedto be within the scope of the claims if they include structural elementsthat do not differ from the literal language of the claims, or if theyinclude equivalent structural elements with insubstantial differencesfrom the literal languages of the claims.

What is claimed is:
 1. A gas turbine engine comprising a compressorsection; a turbine section comprising a drive turbine, the turbinesection located downstream of the compressor section; a fan mechanicallycoupled to and rotatable with the drive turbine such that the fan isrotatable by the drive turbine at the same rotational speed as the driveturbine, the fan defining a fan pressure ratio and comprising aplurality of fan blades, each fan blade defining a fan tip speed;wherein during operation of the gas turbine engine at a rated speed thefan pressure ratio of the fan is less than 1.5 and the fan tip speed ofeach of the fan blades is greater than 1,250 feet per second.
 2. The gasturbine engine of claim 1, wherein during operation of the gas turbineengine at the rated speed the fan pressure ratio is between 1.15 and1.5.
 3. The gas turbine engine of claim 1, wherein during operation ofthe gas turbine engine at the rated speed the fan pressure ratio isbetween 1.25 and 1.4.
 4. The gas turbine engine of claim 1, whereinduring operation of the gas turbine engine at the rated speed the fantip speed of each of the fan blades is between about 1,350 feet persecond and about 2,200 feet per second.
 5. The gas turbine engine ofclaim 1, wherein during operation of the gas turbine engine at the ratedspeed the fan tip speed of each of the fan blades is greater than about1,450 feet per second.
 6. The gas turbine engine of claim 1, whereinduring operation of the gas turbine engine at the rated speed the fantip speed of each of the fan blades is greater than about 1,550 feet persecond.
 7. The gas turbine engine of claim 1, wherein the drive turbineof the turbine section is a low pressure turbine, and wherein the gasturbine engine further comprises a high pressure turbine locatedupstream of the low pressure turbine.
 8. The gas turbine engine of claim7, further comprising: a compressor section comprising a low pressurecompressor and a high pressure compressor, wherein the low pressurecompressor is driven by the low pressure turbine and the high pressurecompressor is driven by the high pressure turbine.
 9. The gas turbineengine of claim 1, further comprising: a nacelle surrounding and atleast partially enclosing the fan.
 10. The gas turbine engine of claim9, further comprising: an inlet pre-swirl feature located upstream ofthe plurality of fan blades of the fan, the inlet pre-swirl featureattached to or integrated into the nacelle.
 11. A method of operating adirect drive gas turbine engine comprising a turbine section with adrive turbine and a fan driven by the drive turbine, the methodcomprising: rotating the fan of the gas turbine engine with the driveturbine of the turbine section of the gas turbine engine such that thefan rotates at an equal rotational speed as the drive turbine, the fandefines a fan pressure ratio less than 1.5, and a fan blade of the fandefines a fan tip speed greater than 1,250 feet per second.
 12. Themethod of claim 11, wherein rotating the fan of the gas turbine engineof the drive turbine comprises operating the gas turbine engine at arated speed.
 13. The method of claim 11, wherein rotating the fan of thegas turbine engine with the drive turbine comprises rotating the fan ofthe gas turbine engine with the drive turbine such that the fan definesa fan pressure ratio between 1.15 and 1.5.
 14. The method of claim 11,wherein rotating the fan of the gas turbine engine with the driveturbine comprises rotating the fan of the gas turbine engine with thedrive turbine such that the fan defines a fan pressure ratio between1.25 and 1.5.
 15. The method of claim 11, wherein rotating the fan ofthe gas turbine engine with the drive turbine comprises rotating the fanof the gas turbine engine with the drive turbine such that the fan bladeof the fan defines a fan tip speed between about 1,350 feet per secondand about 2,200 feet per second.
 16. The method of claim 11, whereinrotating the fan of the gas turbine engine with the drive turbinecomprises rotating the fan of the gas turbine engine with the driveturbine such that the fan blade of the fan defines a fan tip speedgreater than about 1.450 feet per second.
 17. The method of claim 11,wherein rotating the fan of the gas turbine engine with the driveturbine comprises rotating the fan of the gas turbine engine with thedrive turbine such that the fan blade of the fan defines a fan tip speedgreater than about 1,550 feet per second.
 18. The method of claim 11,further comprising: pre-swirling a flow of air provided to the fan ofthe gas turbine engine during operation of the gas turbine engine. 19.The method of claim 18, wherein pre-swirling the flow of air provided tothe fan of the gas turbine engine comprises pre-swirling the flow of airprovided to the fan of the gas turbine engine using an inlet pre-swirlfeature located upstream of the fan blade of the fan and attached to orintegrated into a nacelle of the gas turbine engine.
 20. The method ofclaim 11, wherein the drive turbine of the turbine section of the directdrive gas turbine engine is a low pressure turbine, wherein the turbinesection of the direct drive gas turbine engine further comprises a highpressure turbine, and wherein the direct drive gas turbine enginefurther comprises a compressor section having a low pressure compressorand a high pressure compressor and a nacelle surrounding and at leastpartially enclosing the fan.